Abstract: A turbomachine includes a plurality of transition ducts disposed in a generally annular array. An outlet of each of the plurality of transition ducts is offset from an inlet along the longitudinal axis and the tangential axis. Each of the plurality of transition ducts further includes an upstream portion and a downstream portion, the upstream portion extending between the inlet and an aft end, the downstream portion extending between a head end and the outlet. The turbomachine further includes a support ring assembly downstream of the plurality of transition ducts along a hot gas path. The turbomachine further includes a plurality of support assemblies directly connecting the upstream portion of at least one transition duct of the plurality of transition ducts to the support ring assembly. Each of the plurality of support assemblies includes a rod extending between the upstream portion and the support ring assembly.

Abstract: An axially staged combustion system includes a primary fuel nozzle, a primary combustion zone defined downstream from the primary fuel nozzle, a conical duct disposed downstream from the primary combustion zone and an integrated exit piece disposed downstream from the conical duct. The conical duct and the integrated exit piece at least partially form a hot gas path of the duct section. A plurality of fuel injectors is oriented radially inwardly with respect to an axial centerline of the duct section and is disposed downstream from the primary combustion zone. Each fuel injector of the plurality of fuel injectors provides for injection of a secondary fuel-air mixture into the hot gas path. The plurality of fuel injectors is distributed along at least one of the conical duct and the integrated exit piece.

Abstract: A fuel nozzle assembly includes a fuel plenum body including a forward plate, an aft plate, an outer band extending between the forward plate and the aft plate and a fuel plenum defined within the fuel plenum body. A plurality of tubes extends through the forward wall, the fuel plenum and the aft wall and each tube defines a premix flow passage through and downstream from the fuel plenum body. Each tube is rigidly connected to the aft plate. A damping plate is disposed downstream from the aft plate. At least one tube of the plurality of tubes extends through a corresponding tube opening defined by the damping plate. The damping plate includes a plurality of spring members fixedly connected to the damping plate. At least one spring member of the plurality of spring members is engaged with an outer wall of the at least one tube.

Abstract: An integrated combustor and stage one nozzle in a gas turbine includes a combustion chamber that receives premixed fuel and air from at least one fuel nozzle group at separate axial locations. The combustion chamber includes a liner and a transition piece that deliver hot combustion gas to the turbine. The stage one nozzle, the liner and the transition piece are integrated into a single part. At least one of the axial locations of the one or more fuel nozzle groups includes a plurality of small scale mixing devices that concentrate heat release and reduce flame length.

Abstract: A turbomachine includes a plurality of transition ducts and a support ring assembly. The outlet of a transition duct includes an inner flange and an outer flange. The turbomachine includes bore holes defined in the inner flange and outer flange, mating bore holes defined in the support ring assembly, and mechanical fasteners connecting the inner flange and the outer flange to the support ring assembly. A first one of the bore holes or mating bore holes has a first maximum radial gap and a first maximum tangential gap relative to an associated mechanical fastener. A second one of the bore holes or mating bore holes has a second maximum radial gap and a second maximum tangential gap relative to an associated mechanical fastener. The second maximum radial gap is greater than the first maximum radial gap or the second maximum tangential gap is greater than the first maximum tangential gap.

Abstract: A turbomachine includes a plurality of transition ducts disposed in a generally annular array. Each of the plurality of transition ducts includes an inlet, an outlet, and a passage defining an interior and extending between the inlet and the outlet and defining a longitudinal axis, a radial axis, and a tangential axis. The outlet of each of the plurality of transition ducts is offset from the inlet along the longitudinal axis and the tangential axis. The turbomachine includes a support ring assembly downstream of the plurality of transition ducts along a hot gas path, and a plurality of mechanical fasteners connecting at least one transition duct of the plurality of transition ducts to the support ring assembly. The turbomachine includes a late injection assembly providing fluid communication for an injection fluid to flow into the interior downstream of the inlet of at least one transition duct of the plurality of transition ducts.

Abstract: A turbomachine includes a plurality of transition ducts disposed in a generally annular array. An outlet of each of the plurality of transition ducts is offset from an inlet along the longitudinal axis and the tangential axis. Each of the plurality of transition ducts further includes an upstream portion and a downstream portion, the upstream portion extending between the inlet and an aft end, the downstream portion extending between a head end and the outlet. The turbomachine further includes a support ring assembly downstream of the plurality of transition ducts along a hot gas path. The turbomachine further includes a plurality of support assemblies directly connecting the upstream portion of at least one transition duct of the plurality of transition ducts to the support ring assembly. Each of the plurality of support assemblies includes a rod extending between the upstream portion and the support ring assembly.

Abstract: A turbomachine includes a plurality of transition ducts disposed in a generally annular array. Each transition duct includes an inlet, an outlet, and a passage defining an interior and extending between the inlet and the outlet and defining a longitudinal axis, a radial axis, and a tangential axis. The outlet of each transition duct is offset from the inlet along the longitudinal axis and the tangential axis. Each transition duct further includes an upstream portion and a downstream portion. The turbomachine further includes a late injection assembly disposed between the upstream portion and the downstream portion of a transition duct and which provides fluid communication for an injection fluid to flow into the interior downstream of the inlet of the transition duct. The late injection assembly includes a late injection ring.

Abstract: A segmented annular combustion system includes integrated combustor nozzles, each of which has a fuel injection panel disposed radially between an inner liner segment and an outer liner segment. An aft end portion of the fuel injection panel defines a turbine nozzle. One or more hook plates attached to the outer surfaces of the liner segments are used to mount the integrated combustor nozzle to an inner mounting ring or an outer mounting ring. A mounting strut or mounting tenon(s) attached to the liner segments may also be employed.

Abstract: A segmented annular combustion system includes an alternating arrangement of fuel nozzles and integrated combustor nozzles. The fuel nozzles deliver fuel to the primary combustion zones. The integrated combustor nozzles include an inner liner segment, an outer liner segment, and a fuel injection panel extending between the liner segments. The fuel injection panel includes injection outlets on one or both side walls to deliver a combustible mixture to the secondary combustion zones. Each fuel injection panel, which provides a boundary between adjacent primary and secondary combustion zones, includes an aft end that defines a turbine nozzle. The segmented annular combustion system is part of a gas turbine.

Abstract: A turbomachine includes a plurality of transition ducts disposed in a generally annular array and including a first transition duct and a second transition duct. Each of the plurality of transition ducts includes an inlet, an outlet, and a passage extending between the inlet and the outlet and defining a longitudinal axis, a radial axis, and a tangential axis, the outlet of each of the plurality of transition ducts offset from the inlet along the longitudinal axis and the tangential axis. The turbomachine further includes a support ring assembly downstream of the plurality of transition ducts along a hot gas path, and a plurality of mechanical fasteners connecting at least one transition duct of the plurality of transition ducts to the support ring assembly.

Abstract: A system includes a layered structure. The layered structure includes first and second coalesced layers and an intermediate layer disposed between the first and second coalesced layers. The first and second coalesced layers have a higher degree of coalescence than the intermediate layer.

Abstract: A system includes a layered structure. The layered structure includes first and second coalesced layers and an intermediate layer disposed between the first and second coalesced layers. The first and second coalesced layers have a higher degree of coalescence than the intermediate layer.

Abstract: The present application provides a clocked combustor can array for coherence reduction in a gas turbine engine. The clocked combustor can array may include a number of combustor cans positioned in a circumferential array. A first set of the combustor cans may have a first orientation and a second set of the combustor cans may have a second orientation.

Abstract: The present application provides a combustion system for use with a cooling flow. The combustion system may include a head end, an aft end, a transition nozzle extending from the head end to the aft end, and an impingement sleeve surrounding the transition nozzle. The impingement sleeve may define a first cavity in communication with the head end for a first portion of the cooling flow and a second cavity in communication with the aft end for a second portion of the cooling flow. The transition nozzle may include a number of cooling holes thereon in communication with the second portion of the cooling flow.

Abstract: Transition duct assemblies for turbine systems and turbomachines are provided. In one embodiment, a transition duct assembly includes a plurality of transition ducts disposed in a generally annular array and comprising a first transition duct and a second transition duct. Each of the plurality of transition ducts includes an inlet, an outlet, and a passage extending between the inlet and the outlet and defining a longitudinal axis, a radial axis, and a tangential axis. The outlet of each transition duct is offset from the inlet along the longitudinal axis and the tangential axis. The transition duct assembly further includes an aerodynamic structure defined by the passages of the first transition duct and the second transition duct. The aerodynamic structure includes a pressure side, a suction side, and a trailing edge, the trailing edge having a modified aerodynamic contour.

Abstract: A gas turbine includes a compressor and at least one combustor downstream from the compressor. The combustor includes a burner having an inner shroud extending axially along at least a portion of the burner, an outer shroud radially separated from the inner shroud and extending axially along at least a portion of the burner, and a plurality of stator vanes extending radially between the inner shroud and the outer shroud. The stator vanes have an inner end proximate the inner shroud and an outer end proximate the outer shroud. The burner further includes a vortex tip at one of either the inner end or the outer end of the stator vanes. The vortex tip provides a gap between the inner end and the inner shroud or the outer end and the outer shroud, and the vortex tip includes a plurality of fuel ports. The gas turbine further includes a turbine downstream from the combustor.

Abstract: The present application provides a micro-mixer combustion nozzle for mixing a flow of fuel and a flow of air in a gas turbine engine. The micro-mixer combustion nozzle may include a fuel plate with a number of fuel plate apertures and a fuel plate passage in communication with the flow of fuel and an air plate with a number of air plate apertures and an air plate passage in communication with the flow of air. The fuel plate passage and the air plate passage may align to mix in part the flow of fuel and the flow of air.

Abstract: A turbine system is disclosed. The turbine system includes a transition duct having an inlet, an outlet, and a passage extending between the inlet and the outlet and defining a longitudinal axis, a radial axis, and a tangential axis. The outlet of the transition duct is offset from the inlet along the longitudinal axis and the tangential axis. The turbine system further includes a turbine section connected to the transition duct. The turbine section includes a plurality of shroud blocks at least partially defining a hot gas path, a plurality of buckets at least partially disposed in the hot gas path, and a plurality of nozzles at least partially disposed in the hot gas path. At least one of a shroud block, a bucket, or a nozzle includes means for withstanding high temperatures.

Abstract: A fuel nozzle for a gas turbine generally includes a main body having an upstream end axially separated from a downstream end. The main body at least partially defines a fuel supply passage that extends through the upstream end and at least partially through the main body. A fuel distribution manifold is disposed at the downstream end of the main body. The fuel distribution manifold includes a plurality of axially extending passages that extend through the fuel distribution manifold. A plurality of fuel injection ports defines a flow path between the fuel supply passage and each of the plurality of axially extending passages.